The invention relates to a fan for an aircraft engine, particularly a gas turbine aircraft engine. The invention further relates to an aircraft engine.
Aircraft engines consist of, among other things, a fan and a core engine. The core engine comprises at least a compressor, a combustion chamber and at least one turbine. Conventional aircraft engines have a fan with a fan rotor, which is equipped with fan blades. The fan blades form a blade ring and extend radially outwardly from a hub of the fan rotor to an outer flow path wall of a fan flow path. In conventional aircraft engines of this type, the fan rotor is thus equipped with a single row of blades, i.e., it has only one ring of fan blades. In fans of this type, the mass flow that can be achieved by the fan is limited.
To increase the mass flow generated by a fan, it is known to position an auxiliary rotor upstream of the fan rotor. U.S. Pat. No. 6,722,847 B2, for example, discloses a fan of an aircraft engine having a fan rotor and an auxiliary rotor positioned upstream of the fan rotor. The fan blades of the fan rotor extend radially outwardly to the outer flow path wall of the fan flow path. The auxiliary blades of the auxiliary rotor, however, extend radially outwardly from a hub and end at a substantial distance from the outer flow path wall of the fan flow path. In accordance with U.S. Pat. No. 6,722,847 B2, the auxiliary rotor is an integral component of the fan rotor. As a result, the auxiliary blades are rigidly connected to the fan blades, such that, according to this reference, both rotors must rotate at the same speed and in the same direction of rotation.
Based thereon, it is an object of the invention to provide a novel fan for an aircraft engine, particularly a gas turbine aircraft engine, and a novel aircraft engine.
According to the invention, the auxiliary rotor and the fan rotor are designed as separate rotors, such that the auxiliary rotor can be operated at a higher speed than the fan rotor.
According to the present invention, the auxiliary rotor and the fan rotor are designed as two separate rotors, such that the auxiliary rotor can be operated at a higher speed than the fan rotor. This significantly increases the compression ratio in the range of action of the auxiliary rotor and thereby enables an increase in the mass flow generated by the fan. The increased mass flow rate and the increased compression ratio in the hub region result in a lower compression ratio in the outer fan range for a given diameter of the fan flow path and a specific thrust of the aircraft engine. The lower compression ratio is produced with a lower circumferential speed, such that relative Mach numbers become smaller, and fewer losses and thus an improved efficiency can ultimately be achieved. Furthermore, the reduced circumferential speeds and the reduced compression ratios significantly reduce the noise of the aircraft engine. Another advantage is that foreign bodies and impurities can be better centrifuged out of the air drawn in, so that the risk of damage to and erosion in the core engine is minimized.
According to an advantageous further refinement of the invention, a transmission ratio between the auxiliary rotor speed and the fan rotor speed is variable. The auxiliary rotor and the fan rotor can be operated in the same direction of rotation or in opposite directions of rotation.